A FLIGHT INVESTIGATION OF A TERMINAL AREA NAVIGATION

AND GUIDANCE CONCEPT FOR STOL AIRCRAFT

D . W . Smith, F , Neuman, D . M . Watson, and G. H. Hardy

NASA Ames Research Center

INTRODUCTION

Studies have shown (e.g., refs. 1 and 2) t h a t short-haul aircraft may provide an e f fec t ive t ransportat ion system tha t can operate in to c i t y centers and suburban faci l i t ies . design and development of such a short-haul system, a j o i n t DOT/NASA STOL Operating Systems Experiment Program has been i n i t i a t e d . j o in t program, NASA/Ames has developed an experiments program with the overal l objective of providing information tha t w i l l a id i n the choice of terminal area guidance, navigation, and control system concepts f o r short-haul a i r c r a f t , and invest igat ing operational procedures.

To provide the de ta i led data base required for t h e

A s a pa r t of t h i s

In a short-haul t ransportat ion system, various levels of avionics systems capabi l i ty may be needed. gation, guidance, and control of a i r c r a f t operating i n low-density t r a f f i c conditions and r e l a t i v e l y good weather. More complex and cos t ly automated systems may be economically j u s t i f i a b l e f o r operations i n high-density t raff ic conditions and poor weather. The test data obtained i n t h i s program w i l l pro- vide a bas i s f o r t he select ion of system capabi l i ty t o meet operational requirements (e.g., runway requirements, weather minimums, e t c . ) and will also provide means for estimating the system acceptability and system cost.

Simple, low-cost systems may be adequate fo r navi-

A d i g i t a l avionics system referred t o a s STOLAND nas been purchased and in s t a l l ed (without servos) i n the NASA CV-340 twin-engine t ransport a i r c r a f t . Nineteen t e s t f l i g h t s have been made s ince October 1973 t o obtain preliminary STOLAND performance data i n the manual f l i g h t d i r ec to r mode using time- controlled guidance.

SMLAND is a l so in s t a l l ed (with servos) i n the powered-lift Augmentor Wing J e t STOL research a i r c r a f t ( f ig . 1) described i n reference 3 and a DeHavilland DHC-6 Twin Otter STOL a i r c r a f t . ducted i n these a i r c r a f t t o obtain performance data on both simple and sophis- t i ca t ed avionics system concepts and the corresponding STOL operational procedures. t h e more s igni f icant f l i g h t tes t r e s u l t s obtained i n the CV-340 aircraft.

Investigations w i l l soon be con-

This repor t b r i e f l y describes the system concept and presents

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https://ntrs.nasa.gov/search.jsp?R=19760024068 2018-07-13T21:58:14+00:00Z

SYSTEM CONCEPT AND OPERATION

STOLAND is an integrated digital avionics system having a computer of sufficient size, speed, and capability to perform all terminal area navigation, guidance, and control functions, and to automatically control and guide a STOL test vehicle along a curved reference approach flight path. system are the autopilot modes considered standard for commercial’transport aircraft and an autothrottle. This system was built by Sperry Flight Systems to meet stringent performance and environmental requirements. The major com- ponents of the system are a Sperry 1819A general-purpose digital computer and a data adapter that interfaces all the navigation aids, displays, controls, and servo actuators (fig. 2). The navigation aids include VHF omnirange (VOR), distance measuring equipment (DME), tactical air navigation (TACAN) receiver, instrument landing system (ILS), microwave modular instrument landing system described in reference 4 (MODILS), inertial navigation system (INS), and radio altimeter.

Included in the

The system components installed in the cockpit of the aircraft (fig. 3) include the Sperry RD202A horizontal situation indicator (HSI), control wheel, electronic attitude director indicator (EADI), multifunction display (MFD), MFD control panel, mode select panel (MSP), status panel, and data entry panel. During automatic operation, the pilot monitors the system operation through the various cockpit displays. During flight director operation, the pilot uses the same set of displays for guidance information along the reference flight path and to monitor the system. path flown in the CV-340 is shown in figure 4. leg (waypoints 1-10>, a 180° turn to final approach with a So glide slope occuring half way around the turn (waypoints 10-12), and a final straight-in approach (waypoints 12-14).

An illustration of the approach flight It consists of a long inbound

The navigation system used for the approach provides estimates of position and velocity with respect to a runway coordinate system, which has its origin at the glide-slope intercept point (fig. 4 ) . The position and velocity esti- mate are generated using ground navigation aid information blended in a com- plementary filter with inertial information obtained from body-mounted accelerometers and attitude sensors, and air data obtained from a barometric altimeter and an airspeed sensor. from TACAN except when the aircraft is in MODILS coverage after passing point A (fig. 4). The navigation system also estimates wind velocity utiliz- ing air data. aid information,navigation is accomplished by dead reckoning using air data. Upon regaining radio information, the system automatically switches back to the use of radio data. presented in reference 5.

The ground navigation data are obtained

In the event of a momentary loss of ground radio navigation

A detailed description of the navigation system is

The guidance system used for the approach is based on a flight path, stored in the airborne computer, which is specified by waypoints (X,Y,Z coordi- nates) and associated information such as the radius of turn between waypoints and the maximum, minimum, and nominal airspeed between waypoints. The approach guidance is initiated when the aircraft captures the rear extension of the

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straight line between waypoints 8 and 9 (see dotted line, fig. 4). At waypoint 8, controlled time of arrival (4D) guidance is initiated. Slightly before way- point 10, a predictive bank angle command is given, and just before waypoint 11, a constant vertical acceleration maneuver is performed to acquire the So flight-path angle. The short straight-in section (waypoints 12-13) is the last segment using the 4D guidance laws given below. The remaining flight path to flare is flown with similar lateral and longitudinal guidance laws except for the system gains, which are relatively low from waypoints 1 to 13, and are high from waypoint 13 to flare to assure precise path tracking.

For lateral tracking the guidance law is:

- P 4c - KIYerr + K2* + 4

where

'err P cross track velocity

cross track error

equals zero, for a straight line track +P and

for a circular track where

Vg ground speed

R radius of turn

g acceleration due to gravity

For vertical tracking the guidance law is:

K3 0 c = - h Vg err dt

where

- - 'err 'nom - yI (y = flight-path angle)

altitude error err h "I equals -, inertial flight-path angle derived from the navigation

'I vg system

As previously stated, 4D guidance is initiated at waypoint 8 (fig. 4). From this point, the system attempts to arrive at waypoint 13 at a given time.

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Control of a r r i v a l t i m e a t waypoint 13 i s based only on speed control, which is provided by control l ing the t h r o t t l e as a function of an airspeed er ror . EADI. The airspeed command Vc i s defined as the algebraic sum of a prescribed nominal airspeed (Vnom) and an e r ro r t h a t i s proportional t o an aircraft posi- t ion e r ro r (AS):

In the f l i g h t d i rec tor mode, t he airspeed command i s displayed on t h e

C = 'nom - 0.04 AS (m/sec)

where t o a moving ta rge t , which represents the desired aircraft posi t ion. a i r c r a f t a r r ives a t waypoint 8, t he t a rge t and aircraft posi t ions a r e made t o coincide. time it would take t o f l y from waypoint 8 provided the aircraft f l e w t he path exactly a t the nominal airspeed and there was no wind. To account f o r winds, the posi t ion of the moving t a rge t i s recomputed every 10 sec based on t h e latest estimate of wind ve loc i ty and direct ion. assures tha t t h e t a rge t w i l l a r r ive a t waypoint 13 a t t h e nominal a r r i v a l time while moving a t t h e nominal airspeed. approach, the computed posi t ions of the t a rge t would have s tep changes every 10 sec which would r e s u l t i n excessive t h r o t t l e ac t iv i ty . ac t iv i ty , the time r a t e of change i n the value of l imited t o 6.1 m/sec.

AS i s the dis tance along the t rack from t h e estimated aircraft posi t ion As t he

The computed nominal a r r i v a l t i m e a t waypoint 13 i s based on the

This new computed t a rge t posi t ion

If the wind were changing during the

To l i m i t t h e t h r o t t l e AS i n t h e above equation is

RESULTS AND DISCUSSION

As previously noted, t he primary purpose of f l i g h t tests i n the CV-340

The data pre- was t o va l ida te t h e operation of the STOLAND system and t o obtain a preliminary insight i n to the navigation and guidance system performance. sented are from a s e t of 20 simulated IFR (hooded) approaches conducted during the l a t t e r s tages of t h e tests.

For t he CV-340 f l i gh t s , a i r c r a f t posi t ion data were provided by a modified NIKE-HERCULES tracking radar. a minimum mean-square f i l t e r t o obtain a bes t estimate of t he actual aircraft posit ion.

These tracking data were smoothed with

The data presented i n t h i s report are referenced t o a coordinate system whose or ig in is a t t h e MODILS glide-slope in te rcept point (GSIP) on runway 35 a t Crows Landing NALF (see f i g . 4) . The XY plane i s tangent t o the ear th at the or igin; the X ax i s i s pos i t ive i n the d i rec t ion of landing, t he Y ax is i s pos i t ive t o the r igh t , and the H (a l t i tude) ax is i s pos i t ive up. Repre- sentat ive performance of the guidance and navigation systems along a typical approach is discussed, a s well as summary data f o r a l l approaches.

Performance f o r a Typical Approach

The reference f l i g h t path and an example of a typical approach a r e shown i n f igure 5. The top ha l f of the f igure shows the reference path and the

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downrange-crossrange (X vs Y) p lo t of a i r c r a f t posi t ion, and t h e lower pa r t shows the corresponding altitude-downrange (H v s X) p lo t . shown f o r reference. l a t e r a l and ve r t i ca l deviations from the reference path.

The waypoints are The sum of t h e system e r ro r s is represented by t h e

A s shown i n f igure 5 the approach was i n i t i a t e d a t about 520 m a l t i t ude , During the turn about 280 m t o the r igh t , and 30 m above the reference path.

t o f i n a l approach, t he aircraft remained t o the r i g h t of t he path and then acquired the runway center l ine, maintaining t h a t course f o r t he remainder of the approach. The a i r c r a f t remained about 10 t o 30 m above the reference path during the whole approach. The major e r ror p r i o r t o MODILS acquis i t ion can be a t t r i bu ted t o the effect of a TACAN DME bias . The e r rors a t t r i bu tab le t o the navigation and the guidance systems a r e discussed below.

Navigation- Figure 6 presents t he l a t e r a l (cross track) and ve r t i ca l navigation e r rors f o r t he approach shown i n f igure 5, and the envelope of e r rors experienced i n the 20 simulated IFR approaches. the difference between the onboard estimate of t h e a i r c r a f t posi t ion and t h e tracking radar measured posi t ion. combined e f f ec t of e r ro r s due t o ground navaid and airborne receiver s ignal errors , off-nominal atmosphere effects, small e r ro r s i n the ground radar track- ing data, and t h e bas ic navigation system er rors r e su l t i ng from softwarejhard- ware mechanization. f igure 5.

The e r ro r presented i s

The er ror shown i n these t races is the

The waypoints a r e labeled f o r cross reference with

The envelope of l a t e r a l navigation e r rors a t i n i t i a t i o n of t h e approach a t waypoint 8 a r e a s large a s 200 m. than 70 m a t the i n i t i a t i o n of t h e turn a t waypoint 10, where they s t a r t t o incpease again t o values a s la rge a s 150 m. t ha t these navigation e r rors r e s u l t from TACAN e r ro r s i n both range and a z i - muth. MODILS navigation i s i n i t i a t e d . Navigation e r rors then converge smoothly t o less than 15 m after t r ans i t i on t o MODILS i s completed.

These e r ro r s converge t o a maximum less

Examination of t h e da ta ind ica te

A short time a f t e r passing waypoint 10, a t r ans i t i on from TACAN t o

The envelope of the t i m e h i s tory o f the v e r t i c a l navigation e r ro r shows er rors as large as 24 m a t i n i t i a t i o n of the approach a t waypoint 8. ve r t i ca l navigation e r ro r s are always pos i t ive and are probably a r e s u l t of a b ias i n the baro-altimeter. It should be noted tha t the baro-altimeter refer- ence was set p r io r t o each approach based on information radioed from the con- t r o l tower, which gives a correct barometric a l t i t u d e a t the runway level only. After t r ans i t i on t o MODILS and t h e s t a r t of t he descent a t waypoint 11, the baro-altimeter measurement is slowly blended with and replaced by the more accurate MODILS data t o prevent a s tep change i n estimated a l t i t u d e a t the i n i t i a t i o n of glide-slope t racking, t o a constant value of approximately 5 m. time, although it i s speculated t h a t several e r ro r sources could be the cause. For example, a MODILS DME e r ro r of about 60 m could r e s u l t i n the 5-m er ror . I t i s clear t h a t more accurate navigation i s required f o r f i n a l f lare - e.g., a radio altimeter o r a second, more accurate elevation scanner.

The

The ve r t i ca l navigation e r r o r converges This b i a s is unexplained a t t h i s

Guidance- Figure 7 presents t he lateral and ve r t i ca l guidance e r rors f o r

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the approach shown i n f igure 5 and the envelope of e r rors experienced i n t h e

20 simulated IFR approaches. onboard estimate of pos i t i on and t h e reference f l i g h t path. The waypoints are labeled for c ross re ference with figure 5. The envelope of time h i s t o r i e s of t h e lateral guidance e r r o r shows e r r o r s as l a r g e as 400 m a t t h e i n i t i a t i o n o f t h e approach a t waypoint 8; p r i o r t o switching t o MODILS, these e r r o r s con- verge t o smaller values. On switching t o MODILS from TACAN, t h e lateral navi- gation e r r o r decreases while t h e lateral guidance e r r o r increases, reaching a maximum a t about waypoint 11. This increase i n t h e lateral guidance e r r o r r e s u l t s from a TACAN range b i a s e r r o r t h a t causes t h e aircraft t o f l y on t h e r i g h t of t h e reference path from waypoint 8 t o poin t A (see f i g . 5) . Upon switching t o MODILS, which is a more accurate navigation a id , t h e navigation estimate ind ica t e s t h a t t h e aircraft i s f l y i n g t o t h e r i g h t of t h e reference path, thereby generating a lateral guidance e r r o r while t h e navigation e r r o r converges t o a small value. As a r e s u l t of t h e low gain of t h e guidance sys- tem, t h e aircraft is guided slowly t o t h e re ference path. After passing way- poin t 11, t h e lateral navigation and guidance e r r o r s converge t o small values. As shown i n f i g u r e 7, t h e envelope of t h e lateral guidance e r r o r converges t o about +20 m between waypoints 13 and 14 ( i . e . , 1600 m from touchdown). The envelope of v e r t i c a l guidance e r r o r shows e r r o r s as l a rge as 15 m a t t h e i n i t i - a t i o n of t h e approach a t waypoint 8 and i s genera l ly above t h e des i red path. The magnitude of t h e e r r o r represented by t h e envelope remains approximately constant between waypoints 8 and 10. t r a n s i e n t s occur i n t h e v e r t i c a l guidance e r r o r when t h e navigation switches from TACAN t o MODILS and a t approximately waypoint 11 when t h e descent i s i n i t i a t e d . envelope converges t o about +3 m between waypoints 13 and 14 as a r e s u l t of t h e high-gain guidance l a w and high-gain navigation f i l t e r s used during the f i n a l s t r a i g h t - i n approach.

The e r r o r shown i s t h e d i f fe rence between the

A s shown by t h e s o l i d l i n e i n f igu re 7,

The switching t r a n s i e n t decays and t h e v e r t i c a l guidance e r r o r

Summary Performance Data;

Errors Pr ior t o Flare (h z 30.5 m)

Navigation- Figure 8 shows t h e d i f fe rence between t h e a i r c r a f t pos i t i on as measured by ground r ada r and t h e onboard pos i t i on estimate as t h e a i r c r a f t passed through a window posit ioned a t a nominal a l t i t u d e of 30.5 m on a 5' g l i d e slope. (The symbols represent da ta obtained from f l i g h t s on two d i f f e r e n t days.) The da ta show t h a t t h e aircraft was t o t h e l e f t of t h e runway center- l i n e and above t h e g l i d e s lope f o r t h e majority of t h e approaches. For these da ta , t h e vertical mean e r r o r is 2.4 m above t h e reference g l i d e s lope with a la teral mean e r r o r of 1.9 m t o t h e l e f t of cen te r l ine . The 213 e r r o r s about t h e mean are k2.6 m i n a l t i t u d e and k4.2 m i n t h e l a t e r a l d i r ec t ion .

Guidance- Guidance e r r o r s measured a t an a l t i t u d e of 30.5 m a r e presented i n f i g u r e 9. puted by t h e navigation equations. If t h e guidance e r r o r s were zero, t he da t a po in t s would be c lus t e red on t h e estimated glide-slope cen te r l ine which i s t h e o r i g i n of t h e graph. For these da ta , t h e v e r t i c a l mean e r r o r i s 0.8 m below the g l i d e s lope with a lateral mean e r r o r of 0.8 m t o t h e l e f t of center l ine . The 20 v e r t i c a l and lateral e r r o r s about t h e mean a r e k2 .2 m and k6.8 m, respec t ive ly .

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The re ference i n t h i s case i s t h e MODILS 5' g l ide slope as com-

Comparison ,&E Fl igh t Data with CTOL Requirements

The tes t f l i g h t d a t a were compared with FAA Category I1 f i i g h t d i r e c t o r c e r t i f i c a t i o n cr i ter ia f o r CTOL a i r c r a f t t o determine whether t h e navigation system' under inves t iga t ion might be f e a s i b l e f o r a f l i g h t d i r d c t o r landing on a STOL runway i n marginal weather. The FAA cri teria from AC 120-29 state t h a t on t h e loca l i ze r ,

The FAA c r i t e r i a are included i n f i g u r e 9.

"From an a l t i t u d e 300 feet above runway e l eva t ion on t h e approach pa th t o t h e decision a l t i t u d e (100 f e e t ) , t h e f l i g h t d i r e c t o r should cause t h e a i r p l a n e t o t r a c k t o within F25 microamperes (95-percent probabil- i t y ) of t h e ind ica ted course. The performance should be free of sus- ta ined osc i l l a t ions . "

and on t h e g l i d e slope,

"From 700 f e e t a l t i t u d e t o t h e dec is ion a l t i t u d e (100 feet), t h e f l i g h t d i r e c t o r should cause t h e a i rp l ane t o track t h e center of t h e ind ica ted g l i d e slope t o within +75 microamperes o r A 1 2 f e e t , whichever is 'the- l a rge r , without sustained osc i l l a t ions . "

Based on a conventional CTOL runway arrangement, these cr i ter ia would t r a n s l a t e i n t o allowable devia t ions of about F3.7 m (12 f t ) ver t ical and 521 m (69 f t ) l a t e r a l l y f o r a CTOL a i r c r a f t a t a longi tudina l loca t ion defined by the 30.5-m (100-ft) a l t i t u d e poin t on a 2 . 7 O g l i d e slope.

Figure 9 ind ica t e s t h a t t h e 20 e r r o r s measured i n t h e test f l i g h t s are within those prescribed f o r CTOL Category I1 system landing minima (shaded i n f i g . 9 ) . Additional t e s t i n g i s needed t o de f ine t h e performance cri teria f o r STOL a i r c r a f t c e r t i f i c a t i o n f o r Category I1 weather minima. This comparison of t h e test f l i g h t da t a with t h e FAA c r i t e r i a i s no t e n t i r e l y va l id , because t h e landing system, t h e wind environment, t h e g l i d e slope, and o ther parameters . were d i f f e r e n t from those out l ined i n t h e FAA advisory c i r c u l a r . Nevertheless, it gives some measure of t h e system performance.

Speed Control and Longitudinal Guidance

Figure 10 presents t h e longi tudina l guidance e r r o r (AS), t h e commanded airspeed, t h e t r u e airspeed, and t h e ground speed f o r t h e approach shown i n f igu re 5. ( f ig . 5) and t h e boundaries of t h e allowable airspeed commands, designated by the unshaded area, which are based on the a i r c r a f t performance c a p a b i l i t i e s . A comparison of t h e ground speed and t r u e airspeed i n figure 10 ind ica t e s t h e s t rong headwind conditions experienced by t h e a i r c r a f t on t h e f l i g h t path between waypoints 8 and 10. an airspeed above t h e nominal t o meet t h e spec i f i ed a r r i v a l t i m e . As shown, t h e longitudinal e r r o r , AS, increased l i n e a r l y and t h e airspeed command increased above t h e nominal a i r speed f o r t h e first 3000 m of t r ack d is tance . From waypoints 10 t o 11, AS decreased l i n e a r l y a t i t s ra te l i m i t , as t h e air- craft caught up with t h e t a r g e t and t h e commanded a i r speed approached t h e

Also shown are t h e nominal airspeed spec i f i ed f o r t h e reference pa th ,

Under such conditions, t h e a i r c r a f t should f l y a t

0

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nominal. equivalent to a 1.3-sec time error, remained to be corrected at waypoint 13.

In this approach a longitudinal error, AS, of 76 m, which is

Time-of-Arrival Errors at Waypoint 13

Figure 11 is a histogram of the time of arrival errors at waypoint 13 for the simulated instrument (hooded) approaches. For these tests, the mean time- of-arrival error is 3.7 sec (late) with 20 deviation of k3.4 sec. The mean time-of-arrival error obtained during these tests may result from the TACAN range error which caused the actual longitudinal distance flown to be longer than the reference path. performance for all TACAN errors.

Additional data are required to establish the system

It is interesting to note that current manual guidance techniques enable air traffic controllers to deliver CTOL aircraft to the runway within about 515 sec of the predicted arrival time (ref. 6). This capability corresponds to a single runway acceptance rate of about 40 IFR arrivals per hour using cur- rent separation standards. time of arrival guidance system described here it would be possible to increase the runway acceptance rate by about 40 percent (see ref. 6).

Using the improved capability of the automatic

CONCLUSIONS

Results are presented for 20 flight director approaches made during an investigation of a STOL approach and landing concept using the NASA CV-340 air- craft. Results of these limited tests led to the following conclusions:

1. Blended radio/inertial navigation using TACAN and a microwave scanning beam landing guidance system (MODILS) permitted a smooth transition from area navigation (TACAN) t o precision terminal navigation (MODILS).

2. Guidance system (flight director) performance measured at an altitude of 30.5 m was within that prescribed in FAA AC 120-29 f o r Category II CTOL operations on a standard runway.

3. Time of arrival4 at a point about 2 mi from touchdown was about 4 sec k3 sec (20) later than the computed nominal arrival time.

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REFERENCES

1. Anon.: Civil Aviation Research and Development Policy Study - Report. NASA SP-265, 1971. (Also available as DOT TST-10-4.)

2. Anon.: Civil Aviation Research and Development Policy Study - Supporting Papers. NASA SP-266, 1971. (Also available as DOT TST-10-5.)

3. H. G. Quigley; R. C. Innis; and S . Grossmith: A Flight Investigation of the STOL Characteristics of an Augmented Jet Flap STOL Research Aircraft. NASA TM X-62,334, May 1974.

4. Glen D. Adams: Evaluation of STOL Modular Instrument Landing System (MODILS). National Aviation Facilities Experimental Center, Atlantic City, N. J. 08405. FAA, Dept. of Transportation, Report FAARD-72-4, May 1972.

5. Frank Neuman and David N. Warner, Jr.: A STOL Terminal Area Navigation System. NASA TM X-62,348, May 1974.

6. Anon.: Report of Department of Transportation Air Traffic Control Advisory Committee, Dept. of Transportation, Washington, D. C., Dec. 1969.

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Figure 1.- Augmentor wing jet STOL research aircraft.

Figure 2.- STOLAND flight-test system.

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Figure 3.- STOLAND cockpit installation.

DESIRED AIRCRAFT

MODILS ELEVATION

Figure 4 , - Approach flight path.

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0

AIRCRAFT PATH -I 000 REFERENCE PATH LATERAL

DISTANCE, y, m -2000

-3000 500 400

ALTITUDE, 300 H, m 200

I O 0

0 -1000 -2000 -3000 -4000 -5000 LONGITUDINAL DISTANCE, X, m

Figure 5.- Typical flight path.

400 300 1 --- ENVELOPE OF MAXIMUM

EXCURSIONS LATERAL 200 - FLIGHT PATH, fig. 5 ERROR,m

0

- l o o -200 L 30 r

VERTICAL ERROR, m 0 ---------<-

BLEND MODILS WITH ALTIMETER

I I I

0 20 40 60 80 100 120 140 160 180 TIME, sec

Figure 6.- Navigation errors.

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--- ENVELOPE OF MAXIMUM EXCURSIONS

FLIGHT PATH, fig. 5

, -\ //

---- VERTICAL 0

POINTS

0 20 40 60 80 100 I20 140 160 180 TIME, sec

Figure 7.- Guidance errors.

€ 4

6 E W = o

2 e-

AH = 2.4 m 2 c A H = 2.6m

I I I I I

/ 1 I I I

I I I

-6 I I I I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 IO

LATERAL ERROR, AY, m

Figure 8.- Navigation errors at 30.5 m.

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- AH =-0.8m

2cAH = 2.2 m

- AY = -0.8 m 2cAY = 6.8 m

2cr GUIDANCE FAA CTOL E 6 5 4

ar” 2

E O 8

< -2 W

o_ I- -4

> -6 ar W

LATERAL ERROR, AY, m

Figure 9.- Guidance e r r o r s a t 30.5 m.

0 LONGITUDINAL -100

ERROR, AS, m -200

-300

90 80 70

SPEED, 6o V, m/sec

50 40 30 0 1000 2000 3000 4080 5000 6000 7000

LONGITUDINAL DISTANCE, S , m

Figure 10.- Longitudinal guidance.

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I 2 3 4 5 6 7 8 TIME OF ARRIVAL ERROR, A T f , sec

Figure 11.- Time-of-arrival error at waypoint 13.

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